NASA-TN-D-5392-1969 Experimental pressure distributions for a family of blunt bodies at Mach numbers from 2 49 to 4 63 and angles of attack from 0 degrees to 15 degrees《当马赫数为2 49至4.pdf

1、EXPERIMENTA PRESSURE DISTRIBUTIONS FOR A FAMILY OF BLUNT BODIES AT MACH NUMBERS FROM 2.49 TO 4.63 AND ANGLES OF ATTACK FROM Oo TO 15 5 by Robert L. Stullings, Jr., und Dorothy T. Howell Lungler Reseurch Center Langley Stution, Hampton, Vu. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D.

2、 C. AUGUST 1969 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM 19. Security Classif. (of this repart) Unclassified 2. Government Accession No. I 1. Report No. NASA TN D-5392 EXPERIMENTAL PRESSURE DISTRIBUTIONS FOR A FAMILY OF B

3、LUNT BODIES AT MACH NUMBERS FROM 2.49 TO 4.63 AND ANGLES OF ATTACK FROM Oo TO 15 Robert L. Stallings, Jr., and Dorothy T. Howell 4. Title and Subtitle 7. Author(s) XI. Security Classif. (of this page) 21. No. of Pages 22. Price* Unclassified 85 $3.00 9. Performing Organization Name and Address NASA

4、Langley Research Center Hampton, Va. 23365 12. Sponsoring Agency Nome and Address National Aeronautics and Space Administration 3. Recipients Catalog No. 5. Report Date Auoust 1969 6. Performing Organization Code 8. Performing Orgoniration Report N L-6667 IO. Work Unit No. 124-07-17-02-23 _ 11. Cont

5、ract or Grant No. 13. Type of Report and Period Cover 0. The measured pres- a! 0. For these locations, the In general, the effect of angle of attack on the measured pressure distributions shown in figure 4 is as would be expected. This effect with increasing angle of attack consists of an increase i

6、n the magnitudes of the pressures on the windward side of the model ( = 180O) and a corresponding decrease on the leeward side ($ = Oo). The results shown in figure 4 at a= 15 are replotted in figure 5 to show more clearly the effect of model geometry on the pressure distributions. These results are

7、 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-presented with shoulder radius as the varying parameter for a constant value of nose radius. The model geometry symbol notation corresponds to the same values of rc/d for all parts of figure 5. extra

8、polations extending from the subsonic pressure measurements at the last instrumen- tation station to a pressure corresponding to sonic velocity at s/s2 = 1. For each value of rn/d the pressure distributions for all values of rc/d are bounded by the pressure distributions obtained on the hemispherica

9、l model (rc/d = 0.5) and the spherical cap model (rc/d = 0). With decreasing values of rn/d, the pressure distribution of the spherical cap models approach those of the hemispherical model, and thus the overall extent of the corner-radius effect is reduced. This trend is more apparent on the leeward

10、 side of the models ( = Oo) than on the windward side ( = 180O). The dashed curves shown for these models with rc/d = 0 are The effect of Mach number on the pressure distributions of selected models repre- senting the full range of geometrical variables is shown in figure 6 for an angle of attack of

11、 Oo. The local measured pressures have been normalized by the measured value at s/d = 0 for two reasons: (1) to make them directly comparable with normalized pressurc distributions for M, = 10 from reference 2 and (2) to eliminate errors associated with using a computed value of (Pt,2), which is sig

12、nificantly affected by the possible Mach number variations discussed previously. A comparison of the pressure distributions for the Mach number range shown in figure 6 clearly indicates that the Mach number effect at a! = 0 rapidly diminishes with increasing nose bluntness (decreasing rc/d) and Mach

13、 number. The pressure distributions of the present investigation at M = 4.63 are approximately the same as those shown for M, = 10 from reference 2. The effect of Mach number on pressure distributions for a!= 15 is shown in a 0 because of the limited instrumentation at this COI figure 7 for the Mach

14、 number range of this investigation. A pitot pressure was not measured on most models for dition. In order to minimize the errors associated with using a computed value of p as discussed in the preceding paragraph, the ratio pL/pt,2 was multiplied by the ratio (pt, 2/p,ax)cu,00; the latter ratio eff

15、ectively accounts for longitudinal Mach number gra- dient in the tunnel. In general, increasing Mach number results in a decrease in the magnitude of the pressure distributions; this trend is the same as that shown for a= 0. As would be expected, the Mach number effects on the pressure distributions

16、 of the hemispherical model relative to the respective stagnation points are very similar to the results shown at a!= 0 relative to s/d = 0. For the remaining models, the extent of the Mach number effect is different for the windward and leeward sides. On the windwarc side of the models, the Mach nu

17、mber effect decreases very rapidly with decreasing shoul- der radius for all values of rn/d. This same trend is shown on the leeward sides for ward surface is relatively insensitive to shoulder radius. The pressure distributions t, 2 I I Y rn/d = 00 and 1.933. For smaller values of rn/d, the Mach nu

18、mber effect on the lee- 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-shown for M, = 4.63 of this investigation for a! Oo are believed also to be applicable at higher Mach number flow for the following reasons: effect as shown in figure 7 occurs

19、for the hemispherical model and (2) the pressure dis- tributions on the hemispherical model at a! Oo are very similar to results obtained at CY= 0 as discussed previously. The pressure distributions at a!= Oo are essentially the same as the pressure distribution obtained at M, = 10 (ref. 2). (1) the

20、 maximum Mach number Circumferential Pressure Distributions Circumferential pressure distributions are shown in figure 8 for a selected num- ber of models representing the range of geometrical variables at a location 3 inches (76.2 mm) from the geometric stagnation point and at an angle of attack of

21、 15. Results were not obtained for model 1 at this location for a! 0. Analytical distributions are also shown in figure 8 from the following empirical equation based on the experimental data: pz w - L Pl L (0.072 cos2 - 0.5 cos + 0.428) + A pz L - pt, 2 = ( pt,2 4,2) Pt, 2 where p l,w and plYL are t

22、he measured windward ( = 180) and leeward ( = Oo) pressures, respectively, at a given value of s/d. The analytical distributions are in excellent agreement with the experimental data for the complete range of geometrical variables shown. The analytical distributions are also in excellent agreement w

23、ith the measured distributions for the test range of angle of attack as shown in figure 9 and for the test range of Mach number as shown in figure 10. SUMMARY OF RESULTS Pressure distributions were experimentally determined for a family of blunt bodies at Mach numbers from 2.49 to 4.63 and angles of

24、 attack from 0 to 15. The family con- sisted of bodies of revolution having variable nose and shoulder radii and cylindrical afterbodies 7.5 inches (191 mm) in diameter. The model geometry ranged from a hemi- sphere cylinder to a flat-face cylinder. The results are summarized as follows: 1. At an an

25、gle of attack of Oo, the Mach number effect on the nondimensional mea- sured pressure distributions decreased with increasing Mach number and nose bluntness, The nondimensional pressure distributions at a free-stream Mach number of 4.63 of this investigation were in good agreement with previously pu

26、blished results for a limited number of models at Mach numbers up to 10. 2. At an angle of attack of 15O, the trend of the variation of the nondimensional pressure distributions with nose bluntness and Mach number was similar to that shown 7 Provided by IHSNot for ResaleNo reproduction or networking

27、 permitted without license from IHS-,-,-at an angle of attack of Oo. The results should therefore be applicable for Mach numbers greater than the maximum Mach number (4.63) of this investigation. 3. Circumferential pressure distributions were in good agreement with an empiri- cally derived equation

28、for the test range of model geometry, Mach number, and angle of attack. 4. A comprehensive presentation of the data in tabular and figure form is included for sufficiently small intervals of nose and shoulder radii to enable pressure distributions to be determined - either directly or by interpolati

29、on - for any body of the general shape described. Langley Research Center, National Aeronautics and Space Administration, Langley Station, Hampton, Va., May 26, 1969, 124-07-17-02-23. 8 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EFERENCES 1. Sta

30、llings, Robert L., Jr.: Experimentally Determined Local Flow Properties and Drag Coefficients for a Family of Blunt Bodies at Mach Numbers From 2.49 to 4.63. NASA TR R-274, 1967. 2. Trimmer, L. L.: Study of the Blunt Body Stagnation Point Velocity Gradient in Hyper- sonic Flow. AEDC TR-6899, DDC No.

31、 AD 669378, U.S. Air Force, May 1968. 3. South, Jerry C., Jr.: Calculation of Axisymmetric Supersonic Flow Past Blunt Bodies With Sonic Corners, Including a Program Description and Listing. NASA TN D-4563, 1968. 4. Love, E. S.; Woods, W. C.; Rainey, R. W.; and Ashly, G. C., Jr.: Some Topics in Hyper

32、sonic Body Shaping. AIAA Seventh Aerospace Sciences Meeting, Jan. 1969. 5. Belotserkovskii, 0. M.: Flow With a Detached Shock Wave About a Symmetrical Pro- file. J. Appl. Math. Mech., vol. 22, no. 2, 1958, pp. 279-296. 6. Traugott, Stephen C.: An Approximate Solution of the Direct Supersonic Blunt-B

33、ody Problem for Arbitrary Axisymmetric Shapes. May 1960, pp. 361-370. J. Aerosp. Sci., vol. 27, no. 5, 7. Boison, J. Christopher; and Curtiss, Howard A.: An Experimental Investigation of Blunt Body Stagnation Point Velocity Gradient. A.R.S. J., vol. 29, no. 2, Feb. 1959, pp. 130-135. 8. Van Dyke, Mi

34、lton D.; and Gordon, Helen D.: Supersonic Flow Past a Family of Blunt Axisymmetric Bodies. NASA TR R-1, 1959. 9. Cooper, Morton; and Mayo, Edward E.: Measurements of Local Heat Transfer and Pressure on Six 2-Inch-Diameter Blunt Bodies at a Mach Number of 4.95 and at Reynolds Numbers per Foot Up to 8

35、1 X lo6. NASA MEMO 1-3-59L, 1959. 10. Taylor, Nancy; Hodge, Ward F.; and Burbank, Paige B.: Measurements of a 1/7-Scale Model of a Mercury Capsule at Angles of Attack From Oo to *20 at Mach Numbers of 3.50 and 4.44. Heat-Transfer and Pressure NASA TM X-522, 1961. 11. Stallings, Robert L., Jr.; and H

36、aggard, Kenneth V.: Heat-Transfer and Pressure Measurements on a Preliminary Project Fire Model at Mach 3.51. NASA TN D-2000, 1964. 12. Jones, Robert A.: Heat-Transfer and Pressure Distributions on a Flat-Face Rounded- Corner Body of Revolution With and Without a Flap at a Mach Number of 8. NASA TM

37、X-703, 1962. 9 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-13. Price, Earl A., Jr.; Stallings, Robert L., Jr.; and Howard, Paul W.: Pressure and Heat-Transfer Distributions of 0.1-Scale Gemini Exit and Reentry Models at Mach Numbers of 3.51 and 4

38、.44. NASA TM X-1149, 1965. 14. Lawson, Warren A.; McDearmon, R. W.; and Rainey, R. W.: Investigation of the Pressure Distributions on Reentry Nose Shapes at a Mach Number of 3.55. NASA TM X-244, 1960. 15. Anon.: Manual for Users of the Unitary Plan Wind Tunnel Facilities of the National Advisory Com

39、mittee for Aeronautics. NACA, 1956. 10 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Orifice 1 2* 3 4* 5 6* 7 8* 9 IO* 11 1 z* 13 14* 15 16b 17 18 I? 20 21 22 23 24 2 5f 26“ 21* 2 9* 29* 30* 31* 32“ 33 34 35 36 37 38 39; 40* 41; 429 43; 44t 45. 46“

40、 47t 48“ 49 50 51 52 53 54 55 50t 57. 58 596 601 61* 62* 63t 64: 058 668 67* 68* 691 70* 7 I* 72* 13; 74; TABLE 1.- TABULATION OF PRESSURE MEASUREMENTS AT Ma = 2.49 (a) Model 1 , deg 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 9 0 0 0 45 45 45 45 45 L5 45 45 45 45 45 45 45 45 90 00 91) 90 90 90 90 90 90

41、 90 90 90 90 90 90 90 91) 180 180 1RO 180 270 270 210 210 210 180 180 180 180 180 210 270 210 210 270 s Id .oooo .0267 .0533 .0800 .IO67 .I333 .I600 .le67 .2133 .2400 .2667 .2933 .3200 .34b7 -3733 .4000 .42b7 .4533 .*e00 .5600 .6133 .6667 .1200 -1733 .0400 .0800 .I333 .I861 .2400 .2933 .3461 .LOO0 .

42、4533 .5600 -6133 . b661 .1200 .7133 .O267 .0533 .OB00 .lo61 .1333 .I867 .2400 .2?33 .3461 .4000 .4400 .5600 .6133 .6661 .7200 .1133 .0267 .0533 .0800 . 1 Ob1 .I333 .O261 .0533 .0800 .IO61 .1333 .I867 -2400 -2933 .3467 .4000 -1867 .2400 .2933 .3467 .4000 -15O .9714 .?779 .?a68 .9880 .?44 -9969 -9956

43、.?880 .96?0 .?424 .e829 -0570 .0532 .I380 .I123 .1913 .9070 .0583 .0506 .0545 .0988 .1432 .e313 .0532 .044 3 .0431 .0431 .2014 .9bl8 .9525 -100 .?e50 .9894 .9944 .9932 .9?57 .?932 .?a81 .?754 .9450 .9147 .E415 .0418 .0456 .0786 .I178 -0506 .a957 .051? .Or18 .or18 .0506 . Ob84 .8488 .0455 .0368 .0368

44、 .0355 .I875 .?SI8 .96?1 Measured by I-psi (6895-NIm 2 ) gage. - 50 .9?06 e9942 .9?42 .?916 .?03 -9853 .9752 -9562 .?I94 .E853 .a143 .0368 .026t -0266 .0355 .0506 . 8116 .0368 .OZ6b .I3278 .0291 .0300 .R548 .9368 .02bt .0253 .0253 .1R3t .9903 .984C at a of - 00 .?570 .9574 .?64 .9St4 .?551 .?539 .?E

45、75 .?a70 .?E12 .?7?5 -9736 .?677 .?599 -9508 .0319 .?203 .e941 .0514 .7855 .0329 .0221 .0221 .0227 .0240 .9581 .9?64 .?540 .9886 .?SO4 .9bS8 .9518 .9232 .a612 .0329 .0227 ,0227 .0221 .0240 .9910 .9ste .?959 .?949 .9925 .sea2 ,9792 .?682 .?523 .?208 .e581 .0329 .0221 .0221 .0221 .le51 .99t4 .?958 .9$

46、41 .?538 .9520 .?Sl.! .9564 .55t2 .9553 .S529 .91?3 .9677 ,9519 .?209 .?et6 .97?8 .?be6 .9525 .?I69 .9ebr 50 .9?52 .98?5 .9032 .9743 .?668 -9528 .933? .9073 .0630 .e238 .7517 .03?2 .0291 .02?1 .0291 .0278 .a339 .33b1 .0253 .a253 .0240 .0240 .a554 .0361 .0278 .3266 .02b6 .1060 .9946 .9?58 100 .?a43 .

47、?757 .?682 .9554 .?453 .?289 .?0b6 .a795 .e326 .7?33 .7236 .0519 .0443 .0431 .0431 .0418 .8110 .0481 .0393 .0393 .0380 .0368 .e491 .0469 .03r0 .0380 .0368 .1a53 .ve12 .9935 150 - .9692 -9579 . 9465 .9313 .?1e6 .?0d9 .8781 .E464 .79?5 .1628 .6?69 .04R1 .0418 .0406 .0406 .3406 .1e30 .0469 .0380 .3300

48、.036rl .0368 .Q315 .0481 .03?3 .0393 .1393 .1812 .9710 .9823 11 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TABLE 1.- TABULATION OF PRESSURE MEASUREMENTS AT Mm = 2.49 - Continued (b) Model 2 150 .9672 .961- .955 -942 .907 .879 -842 .813 .765 .b65

49、8 .442 .252 .147 .056 .055 .054. .6941 .051 .9651 .9651 .9631 .890c .780 .094 .0581 .0771 .2021 .a869 .1769 .9715 .9766 .9829 .9854 .9412 .e856 .1769 .9665 .9b65 .9b27 .a881 .7795 .6948 .0569 - Irific - 1 2 3 48 5 6; 73 8* 9 IO* 11 I2f 13 14 15 15 17 18 19 20 21 22 23* 2 4* 25 26 27 28 29“ 30 31+ 32* 334 34: 35 36 37 38 39